Reducing low flight mach number fuel consumption

ABSTRACT

A gas turbine engine for an aircraft, comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine, the first electric machine having a first maximum output power; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine, the second electric machine having a second maximum output power; and an engine controller configured to identify a condition to the effect that the LP turbine is operating in an unchoked regime, and, in response to an electrical power demand being between zero and the first maximum output power, only extracting electrical power from the first electric machine to meet the electrical power demand.

CROSS-REFERENCE TO RELATED APPLICATIONS

This specification is based upon and claims the benefit of priority fromUK Patent Application Number 1908373.2 filed on 12 Jun. 2019, the entirecontents of which are incorporated herein by reference.

TECHNICAL FIELD

This disclosure relates to gas turbine engines.

BACKGROUND

Gas turbine engines featuring electric machines operable as both motorsand generators are known, such as those used for more electric aircraft.Whilst such engines may include a plurality of such electric machinesfor redundancy, they are only coupled to one of the spools. For example,one known configuration includes such electric machines coupled to thehigh-pressure spool of a twin-spool turbofan. Another includes suchelectric machines coupled to the intermediate-pressure spool of atriple-spool turbofan.

An issue with such a configuration is that for a given electrical powerdemand, there is no choice but to supply it from the single spool in theengine. Thus the design of the turbomachinery must be capable ofaccommodating all possible electrical power demands throughout theoperational envelope, which inevitably leads to compromise.

It has therefore been proposed to include an electric machine on two ormore shafts of a multi-spool engine. Whilst numerous documents putforward candidates for the optimal physical implementations of such anarchitecture, few make reference to the optimal control strategy tooperate such configurations.

SUMMARY

In an aspect, there is provided a gas turbine engine for an aircraft,comprising:

a high-pressure (HP) spool comprising an HP compressor and a firstelectric machine driven by an HP turbine, the first electric machinehaving a first maximum output power;

a low-pressure (LP) spool comprising an LP compressor and a secondelectric machine driven by an LP turbine, the second electric machinehaving a second maximum output power; and

an engine controller configured to identify a condition to the effectthat the LP turbine is operating in an unchoked regime, and, in responseto an electrical power demand being between zero and the first maximumoutput power, only extracting electrical power from the first electricmachine to meet the electrical power demand.

In another aspect, there is provided a method of reducing fuelconsumption at low inlet Mach numbers in a gas turbine engine of thetype having a high-pressure (HP) spool comprising an HP compressor and afirst electric machine driven by an HP turbine, the first electricmachine having a first maximum output power, and a low-pressure (LP)spool comprising an LP compressor and a second electric machine drivenby an LP turbine, the second electric machine having a second maximumoutput power, the method comprising:

identifying a condition to the effect that the LP turbine is operatingin an unchoked regime; and

in response to an electrical power demand being between zero and thefirst maximum output power, only extracting electrical power from thefirst electric machine to meet the electrical power demand

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only with referenceto the accompanying drawings, which are purely schematic and not toscale, and in which:

FIG. 1 shows a general arrangement of an engine for an aircraft;

FIG. 2 is a block diagram of the engine of FIG. 1;

FIG. 3 is a block diagram of the interface of the electronic enginecontroller and other systems on the engine of FIG. 1;

FIG. 4 is block diagram of the functional modules of the powercontroller in the electronic engine controller of FIG. 3;

FIG. 5 is block diagram of the functional modules of the classifier inthe power controller of FIG. 4;

FIG. 6 shows a procedure to optimise operation of the engine of FIG. 1during maximum climb or maximum take-off conditions;

FIG. 7 shows a procedure to optimise operation of the engine of FIG. 1in conditions in which the low pressure turbine is unchoked;

FIG. 8 shows a procedure to optimise operation of the engine of FIG. 1during an approach idle condition;

FIG. 9 shows a characteristic for an axial flow compressor;

FIGS. 10A and 10B show transient working lines for, respectively, ahigh-pressure compressor and a low-pressure compressor;

FIGS. 11A and 11B show transient working lines for, respectively, ahigh-pressure compressor and a low-pressure compressor where control ofpower offtake and/or input is implemented;

FIG. 12 shows a procedure to optimise operation of the engine of FIG. 1during an acceleration event;

FIG. 13 shows a plot of fuel-air ratio against mass flow in a combustorfor different operating conditions;

FIG. 14 shows a procedure to optimise operation of the engine of FIG. 1during a deceleration event;

FIGS. 15A and 15B show, respectively, an increase in electrical powerdemand, and the corresponding transient working line for the compressoron the same spool as the motor-generator meeting said power demand;

FIGS. 16A and 16B show, respectively, an increase in electrical powerdemand, and the corresponding transient working line for the compressoron the same spool as the motor-generator meeting said power demand withassistance from an energy storage system;

FIG. 17 shows a procedure to optimise operation of the engine of FIG. 1in the event of an increase in electrical power demand;

FIGS. 18A and 18B show characteristics for, respectively, ahigh-pressure compressor and a low-pressure compressor and the movementof operating point when shaft power is transferred from the low-pressurespool to the high-pressure spool;

FIG. 19 shows a procedure to optimise power offtake and/or transfer toincrease surge margin;

FIGS. 20A and 20B show characteristics for, respectively, ahigh-pressure compressor and a low-pressure compressor and the movementof operating point when shaft power is transferred from thehigh-pressure spool to the low-pressure spool;

FIG. 21 shows a procedure to optimise power offtake and/or transfer toincrease compression efficiency; and

FIG. 22 shows a procedure to implement a speed limiter function.

DETAILED DESCRIPTION

The present invention is described in the context of two-spool, gearedturbofan engine architectures. However, it will be apparent to thoseskilled in the art that the principles of the present invention may beapplied to other engine types including gas turbines with two or morespools, such as direct-drive turbofans, turboprops, or open rotorengines.

FIGS. 1 & 2

A general arrangement of an engine 101 for an aircraft is shown in FIG.1, with an equivalent block diagram of the main components thereof beingpresented in FIG. 2.

In the present embodiment, the engine 101 is a turbofan, and thuscomprises a ducted fan 102 that receives intake air A and generates twoairflows: a bypass flow B which passes axially through a bypass duct 103and a core flow C which enters a core gas turbine.

The core gas turbine comprises, in axial flow series, a low-pressurecompressor 104, a high-pressure compressor 105, a combustor 106, ahigh-pressure turbine 107, and a low-pressure turbine 108.

In use, the core flow C is compressed by the low-pressure compressor 104and is then directed into the high-pressure compressor 105 where furthercompression takes place. The compressed air exhausted from thehigh-pressure compressor 105 is directed into the combustor 106 where itis mixed with fuel and the mixture is combusted. The resultant hotcombustion products then expand through, and thereby drive, thehigh-pressure turbine 107 and in turn the low-pressure turbine 108before being exhausted to provide a small proportion of the overallthrust.

The high-pressure turbine 107 drives the high-pressure compressor 105via an interconnecting shaft 109. The low-pressure turbine 108 drivesthe low-pressure compressor 104 via an interconnecting shaft 110.Together, the high-pressure compressor 105, interconnecting shaft 109and high-pressure turbine 107 form part of a high-pressure spool of theengine 101. Similarly, the low-pressure compressor 104, interconnectingshaft 110 and low-pressure turbine 108 form part of a low-pressure spoolof the engine 101.

The fan 102 is driven by the low-pressure turbine 101 via a reductiongearbox in the form of a planetary-configuration epicyclic gearbox 111.Thus in addition to the low-pressure compressor 104, the interconnectingshaft 110 is also connected with a sun gear 112 of the gearbox 111. Thesun gear 112 is meshed with a plurality of planet gears 113 located in arotating carrier 114, which planet gears 113 are in turn are meshed witha static ring gear 115. The rotating carrier 114 is connected with thefan 102 via a fan shaft 116.

It will be appreciated however that a different number of planet gearsmay be provided, for example three planet gears, or six, or any othersuitable number. Further, it will be appreciated that in alternativeembodiments a star-configuration epicyclic gearbox may be used instead.

In order to facilitate electrical generation by the engine 101, a firstelectric machine 117 capable of operating both as a motor and generator(hereinafter, “HP motor-generator”) forms part of the high-pressurespool and is thus connected with the interconnecting shaft 109 toreceive drive therefrom. In the present embodiment, this is implementedusing a tower-shaft arrangement of the known type. In alternativeembodiments, the HP motor-generator 117 may be mounted coaxially withthe turbomachinery in the engine 101. For example, the HPmotor-generator 117 may be mounted axially in line with the duct 118between the low- and high-pressure compressors.

Similarly, a second electric machine 119 capable of operating both as amotor and generator (hereinafter, “LP motor-generator”) forms part ofthe low-pressure spool and is thus connected with the interconnectingshaft 110 to receive drive therefrom. In the present embodiment, the LPmotor-generator 119 is mounted in the tailcone 120 of the engine 101coaxially with the turbomachinery. In alternative embodiments, the LPmotor-generator 119 may be located axially in line with low-pressurecompressor 104, which may adopt a bladed disc or drum configuration toprovide space for the LP motor-generator 119.

It will of course be appreciated by those skilled in the art that anysuitable location for the HP and LP motor-generators may be adopted.

In the present embodiment, the HP and LP motor-generators arepermanent-magnet type motor-generators. Thus, the rotors of the machinescomprise permanent-magnets for generation of magnetic fields forinteraction with the stator windings. Extraction of power from, orapplication of power to the windings is performed by a power electronicsmodule (PEM) 121. In the present embodiment, the PEM 121 is mounted onthe fancase 122 of the engine 101, but it will be appreciated that itmay be mounted elsewhere such as on the core gas turbine, or in thevehicle to which the engine 101 is attached, for example.

Control of the PEM 121 and of the HP and LP motor-generator is in thepresent example performed by an electronic engine controller (EEC) 123.In the present embodiment the EEC 123 is a full-authority digital enginecontroller (FADEC), the configuration of which will be known andunderstood by those skilled in the art. It therefore controls allaspects of the engine 101, i.e. both of the core gas turbine and themotor-generators 117 and 119. In this way, the EEC 123 may holisticallyrespond to both thrust demand and electrical power demand.

An embodiment of the overall system will be described with reference toFIG. 3, and the control software architecture will be described withreference to FIGS. 4 and 5. The various control strategies implementedin response to various engine operational phenomena will be describedwith reference to FIGS. 6 to 22.

Various embodiments of the engine 101 may include one or more of thefollowing features.

It will be appreciated that instead of being a turbofan having a ductedfan arrangement, the engine 101 may instead be a turboprop comprising apropeller for producing thrust.

The low- and high-pressure compressors 104 and 105 may comprise anynumber of stages, for example multiple stages. Each stage may comprise arow of rotor blades and a row of stator vanes, which may be variablestator vanes (in that their angle of incidence may be variable). Inaddition to, or in place of, axial stages, the low-or high-pressurecompressors 104 and 105 may comprise centrifugal compression stages.

The low- and high-pressure turbines 107 and 108 may also comprise anynumber of stages.

The fan 102 may have any desired number of fan blades, for example 16,18, 20, or 22 fan blades.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0 percent spanposition, to a tip at a 100 percent span position. The ratio of theradius of the fan blade at the hub to the radius of the fan blade at thetip—the hub-tip ratio—may be less than (or on the order of) any of: 0.4,0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28,0.27, 0.26, or 0.25. The hub-tip ratio may be in an inclusive rangebounded by any two of the aforesaid values (i.e. the values may formupper or lower bounds). The hub-tip ratio may both be measured at theleading edge (or axially forwardmost) part of the blade. The hub-tipratio refers, of course, to the gas-washed portion of the fan blade,i.e. the portion radially outside any platform.

The radius of the fan 102 may be measured between the engine centrelineand the tip of a fan blade at its leading edge. The fan diameter may begreater than (or on the order of) any of: 2.5 metres (around 100inches), 2.6 metres, 2.7 metres (around 105 inches), 2.8 metres (around110 inches), 2.9 metres (around 115 inches), 3 metres (around 120inches), 3.1 metres (around 122 inches), 3.2 metres (around 125 inches),3.3 metres (around 130 inches), 340 cm (around 135 inches), 3.5 metres(around 138 inches), 3.6 metres (around 140 inches), 3.7 metres (around145 inches), 3.8 metres (around 150 inches) or 3.9 metres (around 155inches). The fan diameter may be in an inclusive range bounded by anytwo of the aforesaid values (i.e. the values may form upper or lowerbounds).

The rotational speed of the fan 102 may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan 102 at cruise conditions for an engine having a fan diameter inthe range of from 2.5 metres to 3 metres (for example 2.5 metres to 2.8metres) may be in the range of from 1700 rpm to 2500 rpm, for example inthe range of from 1800 rpm to 2300 rpm, or, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 3.2 metres to 3.8metres may be in the range of from 1200 rpm to 2000 rpm, for example inthe range of from 1300 rpm to 1800 rpm, for example in the range of from1400 rpm to 1600 rpm.

In use of the engine 101, the fan 102 (with its associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades on the flow results in an enthalpy rise dH of the flow. A fan tiploading may be defined as dH/U_(tip) ², where dH is the enthalpy rise(for example the one dimensional average enthalpy rise) across the fanand U_(tip) is the (translational) velocity of the fan tip, for exampleat the leading edge of the tip (which may be defined as fan tip radiusat leading edge multiplied by angular speed). The fan tip loading atcruise conditions may be greater than (or on the order of) any of: 0.3,0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all unitsin this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be inan inclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

The engine 101 may have any desired bypass ratio, where the bypass ratiois defined as the ratio of the mass flow rate of the flow B through thebypass duct to the mass flow rate of the flow C through the core atcruise conditions. Depending upon the selected configuration, the bypassratio may be greater than (or on the order of) any of the following: 10,10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, or 17.The bypass ratio may be in an inclusive range bounded by any two of theaforesaid values (i.e. the values may form upper or lower bounds). Thebypass duct may be substantially annular. The bypass duct may beradially outside the core engine 103. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of the engine 101 may be defined as the ratioof the stagnation pressure upstream of the fan 102 to the stagnationpressure at the exit of the high-pressure compressor 105 (before entryinto the combustor). By way of non-limitative example, the overallpressure ratio of the engine 101 at cruise may be greater than (or onthe order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75.The overall pressure ratio may be in an inclusive range bounded by anytwo of the aforesaid values (i.e. the values may form upper or lowerbounds).

Specific thrust of the engine 101 may be defined as the net thrust ofthe engine divided by the total mass flow through the engine 101. Atcruise conditions, the specific thrust of the engine 101 may be lessthan (or on the order of) any of the following: 110 Nkg⁻¹s, 105 Nkg⁻¹s,100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s, or 80 Nkg⁻¹s. The specificthrust may be in an inclusive range bounded by any two of the values inthe previous sentence (i.e. the values may form upper or lower bounds).Such engines may be particularly efficient in comparison withconventional gas turbine engines.

The engine 101 may have any desired maximum thrust. For example, theengine 101 may be capable of producing a maximum thrust of at least (oron the order of) any of the following: 160 kilonewtons, 170 kilonewtons,180 kilonewtons, 190 kilonewtons, 200 kilonewtons, 250 kilonewtons, 300kilonewtons, 350 kilonewtons, 400 kilonewtons, 450 kilonewtons, 500kilonewtons, or 550 kilonewtons. The maximum thrust may be in aninclusive range bounded by any two of the aforesaid values (i.e. thevalues may form upper or lower bounds). The thrust referred to above maybe the maximum net thrust at standard atmospheric conditions at sealevel plus 15 degrees Celsius (ambient pressure 101.3 kilopascals,temperature 30 degrees Celsius), with the engine 101 being static.

In use, the temperature of the flow at the entry to the high-pressureturbine 107 may be particularly high. This temperature, which may bereferred to as turbine entry temperature or TET, may be measured at theexit to the combustor 106, for example immediately upstream of the firstturbine vane, which itself may be referred to as a nozzle guide vane. Atcruise, the TET may be at least (or on the order of) any of thefollowing: 1400 kelvin, 1450 kelvin, 1500 kelvin, 1550 kelvin, 1600kelvinor 1650 kelvin. The TET at cruise may be in an inclusive rangebounded by any two of the aforesaid values (i.e. the values may formupper or lower bounds). The maximum TET in use of the engine 101 may be,for example, at least (or on the order of) any of the following: 1700kelvin, 1750 kelvin, 1800 kelvin, 1850 kelvin, 1900 kelvin, 1950kelvinor 2000 kelvin. The maximum TET may be in an inclusive rangebounded by any two of the aforesaid values (i.e. the values may formupper or lower bounds). The maximum TET may occur, for example, at ahigh thrust condition, for example at a maximum take-off (MTO)condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium-based bodywith a titanium leading edge.

The fan 102 may comprise a central hub portion, from which the fanblades may extend, for example in a radial direction. The fan blades maybe attached to the central portion in any desired manner. For example,each fan blade may comprise a fixture which may engage a correspondingslot in the hub. Purely by way of example, such a fixture may be in theform of a dovetail that may slot into and/or engage a corresponding slotin the hub/disc in order to fix the fan blade to the hub. By way offurther example, the fan blades maybe formed integrally with a centralhub portion. Such an arrangement may be a bladed disc or a bladed ring.Any suitable method may be used to manufacture such a bladed disc orbladed ring. For example, at least a part of the fan blades may bemachined from a billet and/or at least part of the fan blades may beattached to the hub/disc by welding, such as linear friction welding.

The engine 101 may be provided with a variable area nozzle (VAN). Such avariable area nozzle may allow the exit area of the bypass duct to bevaried in use. The general principles of the present disclosure mayapply to engines with or without a VAN.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the art asthe “economic mission”) may mean cruise conditions of an aircraft towhich the gas turbine engine is designed to be attached. In this regard,mid-cruise is the point in an aircraft flight cycle at which 50 percentof the total fuel that is burned between top of climb and start ofdescent has been burned (which may be approximated by the midpoint—Suchcruise conditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance-)between top of climb and start of descent. Cruise conditions thus definean operating point of, the gas turbine engine that provides a thrustthat would ensure steady state operation (i.e. maintaining a constantaltitude and constant Mach number) at mid-cruise of an aircraft to whichit is designed to be attached, taking into account the number of enginesprovided to that aircraft. For example where an engine is designed to beattached to an aircraft that has two engines of the same type, at cruiseconditions the engine provides half of the total thrust that would berequired for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

The cruise conditions may correspond to ISA standard atmosphericconditions at an altitude that is in the range of from 10000 to 15000metres, such as from 10000 to 12000 metres, or from 10400 to 11600metres (around 38000 feet), or from 10500 to 11500 metres, or from 10600to 11400 metres, or from 10700 metres (around 35000 feet) to 11300metres, or from 10800 to 11200 metres, or from 10900 to 11100 metres, or11000 metres. The cruise conditions may correspond to standardatmospheric conditions at any given altitude in these ranges.

The forward speed at the cruise condition may be any point in the rangeof from Mach 0.7 to 0.9, for example one of Mach 0.75 to 0.85, Mach 0.76to 0.84, Mach 0.77 to 0.83, Mach 0.78 to 0.82, Mach 0.79 to 0.81, Mach0.8, Mach 0.85, or in the range of from Mach 0.8 to 0.85. Any singlespeed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Thus, for example, the cruise conditions may correspond specifically toa pressure of 23 kilopascals, a temperature of minus 55 degrees Celsius,and a forward Mach number of 0.8.

It will of course be appreciated, however, that the principles of theinvention claimed herein may still be applied to engines having suitabledesign features falling outside of the aforesaid parameter ranges.

FIG. 3

A block diagram of the interface of the EEC 123 and other engine systemsis shown in FIG. 3.

As described previously, in the present embodiment, the EEC 123 iscoupled with the PEM 121 to control the HP motor-generator 117 and theLP motor-generator 119. In this way, power may either be extracted fromor added to each of shafts 109 and 110.

In the present embodiment, the PEM 121 facilitates conversion ofalternating current to and from direct current. This is achieved in thepresent embodiment by employing a respective bidirectional powerconverter for conversion of ac to and from dc. Thus, as shown in theFigure, the PEM 121 comprises a first bidirectional power converter 301connected with the HP motor-generator 117, and a second bidirectionalpower converter 302 connected with the LP motor-generator 119. The dcsides of the power converters 301 and 302 are in the present exampleconnected with each other to facilitate bidirectional power transferbetween the motor-generators 117 and 119.

In an embodiment, both motor-generators and associated bidirectionalpower converters are rated at the same continuous power. In a specificembodiment, they are rated at from 250 kilowatts to 2 megawatts. In amore specific embodiment, they are rated from 300 kilowatts to 1megawatt. In a more specific embodiment, they are rated at 350kilowatts.

In other embodiments, the HP motor-generator 117 and the firstbidirectional power converter 301 are rated at a different continuouspower than the LP motor-generator 119 and the second bidirectional powerconverter 302. In a specific embodiment, they are rated at from 250kilowatts to 2 megawatts. In a more specific embodiment, they are ratedfrom 350 kilowatts to 1 megawatt. In a more specific embodiment, theyare rated at 350 kilowatts.

Those skilled in the art will be familiar with the term “continuouspower,” i.e. a maximum sustainable power output that does not damagecomponents due to over-current, over-voltage, or over-temperature forexample.

In the present example, the dc sides of the power converters 301 and 302are also connected with a dc bus 303.

In the present example the dc bus 303 has connected to it various loads,which may be either located on the engine 101 or on the vehicle instead.Some such as anti-icing systems 304 may be part of the engine, such aselectric nacelle anti-icing systems, or part of the aircraft on whichthe engine 101 is installed, such as electric wing anti-icing systems.

Other loads may be connected with the dc bus 303 and be able to drawpower from and supply power to the bus, such as an energy storage devicein the form of a battery 305. In the present example, control of thecharge/discharge state of the battery 305 is facilitated by a dc-dcconverter 306. Other energy storage devices may be connected to the dcbus 303 as well as or in place of the battery 305, such as a capacitor.

Other loads, indicated at 307, may be connected with the dc bus 303 suchas cabin environmental control systems, electric actuation systems,auxiliary power units, etc.

In operation, the EEC 123 receives a plurality of demand signals, namelya thrust demand in the form of a power lever angle (PLA) signal whichmay be manually set or by an autothrottle system, and an electricalpower demand (P_(D)). In addition, the EEC receives a plurality of setsof measured parameters, namely a set concerning flight parameters of thevehicle, Σ_(AIRCRAFT), and a set concerning operational parameters ofthe engine, Σ_(ENGINE). As will be described further with reference toFIGS. 4 and 5, these demands and parameters facilitate the derivation ofa set of output parameters to control the core gas turbine and themotor-generators.

Thus in addition to the routine set of output parameters generated by aFADEC, such as the fuel flow W_(F) to be metered by a fuel metering unit308 on the engine 101, or variable-stator vane angle, handling bleedsettings, etc. in the present embodiment the EEC 123 comprises a powercontroller module 309 to generate a control signal P_(H) for the firstbidirectional power converter 301 and a second control signal P_(L) forthe second bidirectional power converter 302. The control signals P_(H)and P_(L) control the operation of the power converters in terms of bothdirection and magnitude of electrical power. In this way, the EEC 123may meet the demanded power P_(D) using a suitable balance of electricalpower from the HP and LP motor-generators. As will be described withreference to the later Figures, the optimum way to do this variesthroughout a mission.

In addition to the control signals P_(H) and P_(L), in the presentexample the power controller 309 is configured to derive a controlsignal P_(BAT) for the dc-dc converter 306 to facilitate charge ordischarge of the battery 305.

In a specific embodiment, the power controller 309 is configured to, innormal operation, set P_(BAT) to zero, and only change its status as setout in the optimisation routines described herein, for example theroutines described with reference to FIGS. 8 and 12.

In another specific embodiment, the power controller 309 includesbattery optimisation functionality and modifies the power demandparameter P_(D) by adding or subtracting a value P_(BAT) depending uponwhether it is more optimal to charge, discharge, or maintain the chargeof the battery 305. Those skilled in the art will be familiar with suchtypes of battery state-of-charge optimisation routines.

Thus, in such an embodiment, the power controller 309 modifies the powerdemand parameter P_(D) by performing an addition-assignment operationP_(D)+=P_(BAT). The sign convention used herein is such that a positiveP_(BAT) means that the battery 305 is to be charged, and thus additionalgeneration is required from the HP and LP motor-generators, whilst anegative P_(BAT) means that the battery 305 is to be discharged.

It will be appreciated that in embodiments lacking an energy storagedevice, this signal will not be generated and thus the power demandparameter P_(D) will not be modified at all.

In the present example another control signal P_(AI) is generated toactivate the anti-icing systems 304. In the present example only thenacelle anti-ice system of the engine 101 is controlled, however it isenvisaged that the EEC 123 may in alternative implementations haveauthority over wing anti-ice in certain circumstances. In a similar wayto the modification of P_(D) by the addition-assignment of P_(BAT), thepower controller 309 is configured to perform the addition assignmentP_(D)+=P_(AI) so as to account for the power required for the anti-icingsystems 304. Again, it will be appreciated that in embodiments lackingelectric anti-ice systems, this signal will not be generated and nomodification of P_(D) will be performed in this manner.

FIG. 4

In practice, the EEC 123 houses microprocessors for executing programmodules to control the engine 101. A block diagram of the functionalmodules of the power controller 309 is shown in FIG. 4.

Input parameters previously described with reference to FIG. 3 areinitially received by a classifier module 401 to output an optimisersetting mode for an optimiser module 402. The operation of theclassifier module 401 will be described further with reference to FIG.5, and the various modes of the optimiser module will be described withreference to FIGS. 7 to 22.

The input parameters are also supplied to a filter 403 prior to updatingan engine model module 404. The filter 403 in the present example is anintegrator to smooth out short-term transients so as to not cause largevariations in the engine model. The engine model module 404 runs a realtime model of the engine 101 so as to facilitate prediction of changesis operational parameters, such as W_(F), P_(H), and P_(L) given athrust demand. Such models and their integration within an EEC will befamiliar to those skilled in the art.

Following entry into a given optimisation setting, the optimiser module402 finds the optimal set of parameters for operation of the engine 101given the current operational state of aircraft on which the engine isinstalled and the engine itself.

FIG. 5

The classifier module 401 is detailed in FIG. 5.

In the present embodiment, the classifier module 401 comprises acomparator 501 which compares the present altitude (ALT), the Machnumber (Mn) and the power lever angle to determine the flight regime. Itwill be appreciated that other inputs may be utilised to increase thefidelity of the comparison process, such as engine conditions, ambienttemperature, etc.

In the present embodiment, the comparator is configured to identify ifthe engine is operating in a maximum take-off condition if the altitudeis less than a threshold, the Mach number is less than a threshold, andthe power lever angle is at a maximum. In an embodiment, the altitudethreshold is 5000 feet, whilst the Mach number threshold is 0.3.

In the present embodiment, the comparator is configured to identify ifthe engine is operating in a maximum climb condition if the altitude isabove a threshold, and the power lever angle is at a maximum. In anembodiment, the altitude threshold is 30000 feet.

The optimisation strategy for the maximum take-off condition and themaximum climb mode of operation will be described further with referenceto FIG. 6.

In the present embodiment, the comparator is configured to identify ifthe engine is operating in a cruise if the altitude is above athreshold, the Mach number is in a cruise range, and the power leverangle is at a cruise setting. In an embodiment, the altitude thresholdis 30000 feet, whilst the Mach number range is from 0.8 to 0.9.

In such circumstances, the optimisation strategy that can be employed isto minimise fuel flow W_(F) at constant thrust. Alternatively, theoptimiser may be set to optimise surge margin in the engine, or tooptimise compression efficiency depending on the engine and aircraftparameters. Such strategies will be described further with reference toFIGS. 19 and 21, respectively.

Alternatively, the optimiser may be set to optimise the bypass ratio byvarying the core flow, implementing a variable cycle.

In the present embodiment, the comparator is configured to identify ifthe engine is operating in a regime in which the low-pressure turbine108 is operating in an unchoked condition. This typically manifests atlow Mach number idle conditions, although the unchoked condition mayexist at other operating points depending upon the specific design ofthe low-pressure turbine 108 and the rest of the engine 101. In thepresent embodiment, the comparator is configured to identify thiscondition if the Mach number is below a threshold, and the power leverangle is at an idle setting. In an embodiment, the Mach number thresholdis 0.2. In other embodiments, other inputs such as the corrected speedof the low pressure spool may be utilised to identify the unchokedcondition.

The optimisation strategy for the low Mach number idle condition will bedescribed further with reference to FIG. 7.

In the present embodiment, the comparator is configured to identify ifthe engine is operating in an approach Mach number idle condition if thealtitude is within a range, the Mach number is in a range, and the powerlever angle is at an idle setting. In an embodiment, the altitude rangeis from 0 to 5000 feet above ground level, and the Mach number range isfrom 0.2 to 0.3.

The optimisation strategy for the approach idle condition will bedescribed further with reference to FIG. 8.

The classifier module 401 further comprises a differentiator 502 whichis configured to monitor the PLA and P_(D) parameters and identify atransient type.

In response to the change in power lever angle being positive, thedifferentiator 502 is configured to identify the initiation of anacceleration event. The optimisation strategy for this manoeuvre will bedescribed further with reference to FIG. 12.

In response to the change in power lever angle being negative, thedifferentiator 502 is configured to identify the initiation of adeceleration event. The optimisation strategy for this manoeuvre will bedescribed further with reference to FIG. 14.

In response to a change in electrical power demand P_(D), thedifferentiator 502 is configured to cause the optimiser to invoke theoptimisation strategy described with reference to FIG. 17.

The classifier module 401 further comprises a limiter 503 which isconfigured to monitor the high-pressure and low-pressure spool speeds,N_(H) and N_(L). Should either spool speed approach a limit, which maybe a maximum limit or a keep-out zone, the optimisation strategydescribed with reference to FIG. 22 may be invoked.

In the present example, the outputs of the comparator 501,differentiator 502 and limiter 503 are compared by a prioritiser 504. Itwill be appreciated that there may be concurrent outputs from eachinitial stage of the comparator module, and thus in the presentembodiment the comparator is configured to filter to only one outputoptimiser setting. In the present embodiment, outputs from the limiter503 are priorities over outputs from the differentiator 502, which arein turn prioritised over outputs from the comparator 501.

FIG. 6

Following the identification of a maximum take-off or maximum climbcondition by the classifier module 401, the optimiser 402 enters thecorresponding optimisation routine at step 601. At step 602, a questionis asked as to whether the power demand P_(D) is less than the maximumpower rating of the LP motor-generator 119, P_(Lmax). If so, thencontrol proceeds to step 603 where the optimiser 402 maximises the powergeneration by the LP motor-generator 119, P_(L).

Preferring electrical offtake from the low-pressure spool attemperature-limited conditions such as maximum take-off and maximumclimb conditions reduces the load on the high-pressure spool. For agiven power demand P_(D), this results in a higher high-pressure spoolrotational speed N_(H). This increases the core flow C through the coregas turbine, and so reduces the stator outlet temperature required todeliver the low-pressure turbine power for a given thrust.

It has been found that for motor-generators rated at 350 kilowatts, itis possible to reduce the stator outlet temperature by 2 kelvin. It willbe appreciated that the higher the rating of the motor-generators, thegreater the reduction that may be achieved.

Clearly, if the power demand P_(D) after accounting for batterycharge/discharge and/or anti-icing system operation is greater than themaximum power rating of the LP motor-generator 119, P_(Lmax), then thequestion asked at step 602 will be answered in the negative. In thiscase, control proceeds to step 604 where the optimiser 402 maximises thepower generation by the LP motor-generator 119, P_(L), and minimises thepower generation by the HP motor-generator 117, P_(H). Thus, the LPmotor-generator 119 is directed to generate P_(Lmax) and the HPmotor-generator 117 is directed to generate the remainder,P_(D)-P_(Lmax).

In an embodiment, should spare capacity be available from the LPmotor-generator 119, then the optimiser 402 may elect to divertP_(Lmax)-P_(D) to the HP motor-generator 117 which may further increasecore flow and reduce stator outlet temperature.

FIG. 7

Following the identification of an unchoked regime for the low-pressureturbine 108 by the classifier module 401, the optimiser 402 enters thecorresponding optimisation routine at step 701. As described previously,this may be triggered by a low Mach number idle operating condition, forexample the ground idle operating point.

At step 702, a question is asked as to whether the power demand P_(D) isless than the maximum power rating of the HP motor-generator 117,P_(Hmax). If so, then control proceeds to step 703 where the optimiser402 maximises the power generation by the HP motor-generator 119, P_(H).

As the low-pressure turbine 108 is unchoked, there is a large impactwhen it is required to provide electrical power via the LPmotor-generator 119. The reduction in the speed of the low-pressurespool, N_(L), is such that there is a steep drop in the efficiency ofthe low-pressure compressor 104, and thus an increase in fuel burn. Inpractice, there may also be a drop in efficiency of the LPmotor-generator 119 due to the lower rotational speed. Thus, at the lowflight Mach numbers that cause the low-pressure turbine 108 to unchoke,it is possible to improve fuel consumption by this optimisation routine.

In a similar way to the situation of FIG. 6, should the power demandP_(D) be greater than the maximum power rating of the HP motor-generator117, P_(Hmax), then the question asked at step 702 will be answered inthe negative. In this case, control proceeds to step 704 where theoptimiser 402 maximises the power generation by the HP motor-generator117, P_(H), and minimises the power generation by the LP motor-generator119, P_(H). Thus, the HP motor-generator 117 is directed to generateP_(Hmax) and the LP motor-generator 119 is directed to generate theremainder, P_(D)-P_(Hmax).

In an embodiment, should spare capacity be available from the HPmotor-generator 117, then the optimiser 402 may elect to divertP_(Hmax)-P_(D) to the LP motor-generator 119 which may further reducefuel consumption.

FIG. 8

Following the identification of an approach idle condition by theclassifier module 401, the optimiser 402 enters the correspondingoptimisation routine at step 801. At step 802, a question is asked as towhether the power demand P_(D) is less than the maximum power rating ofthe LP motor-generator 119, P_(Lmax). If so, then control proceeds tostep 603 where the optimiser 402 maximises the power generation by theLP motor-generator 119, P_(L). The excess capacity of the LPmotor-generator 119, P_(Lmax)-P_(D), is transferred to the HPmotor-generator 117.

This has two effects. First, extraction of the maximum power possiblefrom the low-pressure spool significantly reduces the low-pressure spoolrotational speed, N_(L). Recalling that in the engine 101 the fan 102,which is primary thrust generating element, rotates with thelow-pressure spool, it will be clear that a reduction in N_(L) reducesthe thrust generated by the engine 101. This relaxes the requirement touse high drag devices on the airframe to achieve a required descentrate. This reduces noise and reduces fuel consumption.

Second, on approach, the engine idle setting is often constrained by therequirement for the engine to respond to a throttle transient in atimely manner—in the event of a go-around, the engine must deliver fullrated thrust as quickly as possible. Initial high-pressure spool speedN_(H) has a powerful effect on the response time during an engineacceleration, so maximising N_(H) at idle significantly reduces theengine acceleration time. However, this is usually at the expense of alow idle thrust.

The inventor has found that by transferring power from the low-pressurespool to the high-pressure spool allows a higher N_(H) and thus ashorter response time, along with a reduced idle thrust due to the lowerN_(L). It has been demonstrated that in an engine of the type describedherein, 105 kilowatts of power transfer achieves a sufficiently highN_(H) and a 75 percent reduction in idle thrust.

If the power demand P_(D) is greater than the maximum power rating ofthe LP motor-generator 119, P_(Lmax), then the question asked at step802 will be answered in the negative. In this case, control proceeds tostep 804 where the optimiser 402 maximises the power generation by theLP motor-generator 119, P_(L), and minimises the power generation by theHP motor-generator 117, P_(H). Thus, the LP motor-generator 119 isdirected to generate P_(Lmax) and the HP motor-generator 117 is directedto generate the remainder, P_(D)-P_(Lmax).

In an embodiment, the optimiser 402 is further configured to identifythat maintaining the requisite N_(H) will cause unsafe operation of thelow-pressure compressor 104. This may be caused by the operating pointof the LP low-pressure compressor 104 becoming too close to surge or tochoke. In response to the onset of such a condition, the fuel flow W_(F)may be increased.

Alternatively, in order to reduce fuel burn on approach, the optimiser402 may be configured to supplement the HP motor-generator using anenergy storage device, for example the battery 305 via control of theP_(BAT) parameter, or another energy source such as the auxiliary powerunit on the aircraft. In this way, community emissions on approach maybe reduced.

FIG. 9

One of the primary considerations for control of a gas turbine engine isthe prevention of surge in the compression stages. A characteristic 901for an axial flow compressor, such as low-pressure compressor 104 orhigh-pressure compressor 105, is shown in FIG. 9.

The characteristic 901 plots pressure ratio against flow function, whichin this case is non-dimensional flow (W√T/P). The characteristic 901shows a plurality of non-dimensional speed lines 902, 903, 904, 905,along with the compressor steady state working line 906, which is thelocus of operating points for various steady state throttle settings atdifferent non-dimensional speeds. In addition, the surge line 907 isshown, which is the locus of points at which the compressor enters surgeat the various non-dimensional speeds. For a given value of the flowfunction, the pressure ratio at which surge is encountered is denoted R.The difference in pressure ratio on the working line 906 and the valueof R on the surge line 907 for a given value of the flow function isdenoted dR. Therefore, the surge margin for a given compressor operatingpoint may be defined as dR/R.

It is important to maintain a degree of surge margin at all points inthe operational envelope. This is to mitigate random threats, such asinlet flow instabilities due to crosswinds or turbulence, for example.To a first order, it is often recommended for low-pressure compressorsto have around 15 percent surge margin, and high-pressure compressors tohave around 20 percent surge margin. A significant proportion of thesurge margin, typically up to half, is to make allowance for transientexcursions of the working line during acceleration and decelerationmanoeuvres. Such transient phenomena will be described further withreference to FIGS. 10A, 10B, 11A, and 11B.

FIGS. 10A & 10B

Characteristics for the high-pressure compressor 105 and thelow-pressure compressor 106 showing transient phenomena duringacceleration events are plotted in FIGS. 10A and 10B respectively.

In FIG. 10A, the high-pressure compressor steady-state working line 1001is shown along with lines of constant corrected speed 1002, 1003, 1004,and 1005, and the surge line 1006. During an acceleration manoeuvre, thehigh-pressure compressor 105 moves from an initial operating point 1007to a final operating point 1008 via a transient working line 1009 abovethe steady-state working line 1001.

Similarly, in FIG. 10B, the low-pressure compressor steady-state workingline 1011 is shown along with lines of constant corrected speed 1012,1013, 1014, and 1015, and the surge line 1016. During an accelerationmanoeuvre, the low-pressure compressor 106 moves from an initialoperating point 1017 to a final operating point 1018 via a transientworking line 1019 which crosses the steady-state working line 1011.

During the acceleration manoeuvre, the high-pressure compressor 105initially moves towards surge due to the flow compatibility requirementwith the high-pressure turbine 107. The flow function of thehigh-pressure turbine 107 (W₄₀₅√T₄₀₅/P₄₀₅) is substantially fixed duringmost operating conditions of the engine 101, due to the nozzle guidevanes therein being choked. In order to initiate the accelerationmanoeuvre, for example due to an increase in power lever angle setting,the amount of fuel metered by the fuel metering unit 308 (W_(F)) isincreased. This leads to an increase in turbine entry temperature, T₄₀₅.Normally, the high-pressure spool speed N_(H) is prevented from changinginstantaneously due to its inertia. Consequently, the operating point ofthe high-pressure compressor 105 moves up a line of constant correctedspeed. As the overfuelling continues and the high-pressure spool inertiais overcome, the operating point moves along the transient working line1009 parallel to the surge line 1006. As the acceleration finishes, thehigh-pressure compressor 105 adopts its final operating point 1008 onthe steady-state working line 1001.

Referring now to FIG. 10B, at the initiation of the accelerationmanoeuvre, the operating point of the low-pressure compressor 104 movesa little towards surge, and then crosses the steady-state working line1011. The initial move towards surge is due to the reduction in flow inthe high-pressure compressor 105 due to the high-pressure spool inertiaas described above. As the speed of the high-pressure compressor 105increases it can accept more flow. However, due to the greater inertiaof the low-pressure spool (recalling that it drives the fan 102 via thegearbox 111), it cannot accelerate at the same rate and so the operatingpoints of the low-pressure compressor 104 during the accelerationmanoeuvre fall below the steady-state working line 1011.

It will be understood that handling during acceleration manoeuvres maysignificantly affect the design of the compressor turbomachinery andimpose requirements for systems to manage the transients by eithermodifying the transient working line or the surge line, such as variableguide vanes and bleed valves.

FIGS. 11A & 11B

By contrast, in the engine 101 it is possible to utilise the HPmotor-generator 117 and optionally the LP motor-generator 119 to reducethe transient excursion.

FIG. 11A shows the characteristic for the high-pressure compressor 105when the HP motor-generator 117 is used to overcome the high-pressurespool inertia. It can be seen that for the same degree of overfuelling,the transient working line 1101 is much closer to the steady-stateworking line 1001 and further from the surge line 1006. Thus for a givencompressor configuration, this technique may either be used to improvesurge margin during an acceleration manoeuvre, or facilitate a greaterdegree of overfuelling (up to the stator outlet temperature limit) andthus a faster acceleration time.

FIG. 11B shows the characteristic for the low-pressure compressor 104during application of the same technique on the low-pressure spool usingthe LP motor-generator 119. It may be seen that the transient workingline 1102 is again much closer to the steady-state working line 1011 dueto the reduction in effective low-pressure spool inertia by the LPmotor-generator 119.

FIG. 12

Steps carried out by the optimiser 402 to achieve the advantagesdescribed previously for an acceleration event are set out in FIG. 12.

Following the identification of an acceleration condition by theclassifier module 401, the optimiser 402 enters the correspondingoptimisation routine at step 1201. In the present embodiment, with anenergy storage system such as the battery 305 available on the dc bus303, a question is asked at step 1202 as to whether the battery state ofcharge is greater than a minimum value. As will be appreciated by thoseskilled in the art, this may not be absolutely zero but instead will bea minimum state of charge, for example 20 percent, below which thebattery may be damaged.

If so, a further question is asked at step 1203 as to whether theunmodified total aircraft power demand P_(D) (i.e. prior to anymodification thereof to account for battery optimisation) is less thanthe maximum power available from the battery 305, P_(BATmax). If so,then control proceeds to step 1204 where the optimiser 402 overrides anyconcurrent battery optimisation processes, and fully supplies the powerdemand Po using the battery 305, with any excess being supplied to theHP motor-generator 117 to overcome the HP spool inertia. Optionally, anyfurther excess may be supplied to the LP motor-generator 119.

Should either of the questions asked at steps 1202 or 1203 be answeredin the negative, i.e. the battery 305 has a minimal state of charge, orthe unmodified total aircraft power demand P_(D) is greater than themaximum power of the battery 305, P_(BATmax) (or if indeed theparticular embodiment of the engine 101 does not include a battery),then control proceeds to step 1205 in which power generation by the LPmotor-generator 119, P_(L), is maximised and power generation by the HPmotor-generator 117, P_(H), is minimised.

Following optimisation of the power generation strategy to satisfy thepower demand P_(D) in the preceding steps, the fuel flow W_(F) meteredby the fuel metering unit 308 is increased at step 1206.

FIG. 13

When a deceleration event is initiated, fuel flow by the fuel meteringunit 308 is reduced. In the opposite sense to the scenario describedabove, as fuel flow is decreased, the turbine entry temperaturedecreases instantaneously. This is because, as described previously,during the majority of operating scenarios, the nozzle guide vanes inthe high-pressure turbine 107 are choked and the flow function remainsconstant. As the high-pressure spool speed N_(H) is prevented fromchanging instantaneously due to its inertia, the reduction in turbineentry temperature T₄₀₅ causes the operating point of the high-pressurecompressor 105 to move down a line of constant corrected speed on itscharacteristic. This manifests as an increase in mass flow W₃₁ and adecrease in pressure P₃₁ at the exit of the high-pressure compressor105.

The result of this for the combustor 106 is that not only is the amountof fuel delivered lower, but the mass flow W₃₁ therethrough hasincreased. This means that the combustor 106 operates at a lowerfuel-air ratio (FAR) than normal, which risks weak extinction (alsoknown as lean blowout).

Referring to FIG. 13, which is a plot of FAR against flow function, theweak extinction boundary 1301 for the combustor 105 is illustrated.Corrected flow through the combustor 105 to the right of the weakextinction boundary 1301 results in extinction of the flame and is anunacceptable operating condition. A steady-state FAR for a particularmass flow through the combustor 106, is shown at point 1302. Theconstraint on how aggressive a deceleration manoeuver may be is dictatedby the allowable underfuelling margin. In prior art engines, in whichthe flow function increases slightly at the outset of the decelerationmanoeuver, the underfuelling margin is limited to M₁ due to theproximity of the fuel-air ratio to the weak extinction boundary 1301during the deceleration, shown at point 1303.

However, by actively reducing the speed of the high-pressure spool atthe initiation of and during the deceleration event by the HPmotor-generator 107, the operating point of the high-pressure compressor105 is no longer forced to move down a constant speed line at the outsetof the manoeuvre. Instead, there is only a slight deviation from thesteady state working line due to the reduction in turbine entrytemperature, T₄₀₅. As the speed of the high-pressure compressor 105substantially instantaneously begins to decelerate, the mass flowthrough the combustor 106 also reduces substantially instantaneously. Asshown at point 1304 in FIG. 13, the additional margin M₂ allows agreater degree of underfuelling.

It will be understood that this approach allows the design of thecombustor 106 to be optimised due to the greater weak extinction margin,and also allows the vehicle design to be optimised as a greater thrustreduction is achievable by the engine alone without resort to high dragdevices to reduce forward airspeed in, for example, a slam decelmanoeuvre.

It will also be appreciated that the approach provides a method ofcontrolling weak extinction in the combustor 105. Using the engine model404, for example, the onset of weak extinction may be identified byevaluating the current fuel-air ratio in the combustor 105. This may beachieved, for example, by utilising the flight Mach number, altitude andtemperature to determine the mass flow in the engine 101, thecharacteristic of the fan 102 to determine the mass flow C into the coregas turbine, and the characteristics of the compressors 104 and 105 todetermine the mass flow into the combustor 105. This may be combinedwith the commanded fuel flow W_(F) along with a model of the combustionprocess to determine the fuel-air ratio.

In response to identifying that the fuel-air ratio is approaching theweak extinction boundary 1301, the EEC 123 may use the power controller309 to extract mechanical shaft power from the high-pressure spool usingthe HP motor-generator 117 to prevent a further drop in fuel-air ratioin the combustor 105.

FIG. 14

Steps carried out by the optimiser 402 to achieve the advantagesdescribed previously for a deceleration event are set out in FIG. 14.

Following the identification of a deceleration condition by theclassifier module 401, the optimiser 402 enters the correspondingoptimisation routine at step 1401. A question is asked at step 1402 asto whether the power demand P_(D) is less than the maximum powergeneration capability of the HP motor-generator 117, P_(Hmax). If so,then control proceeds to step 1403 whereupon the power generation of theHP motor-generator 117, P_(H), is maximised to satisfy P_(D).

In the present embodiment, the excess capacity P_(Hmax)-P_(D) istransferred to other loads. In an embodiment, the excess capacity isdirected to an energy storage system, such as the battery 305. Asdescribed previously, the energy storage system may additionally oralternatively comprise a capacitor. Additionally or alternatively, theexcess capacity may be directed to an electrical consumer such as theanti-icing system 304, which may be the nacelle anti-ice system of theengine 101. Alternatively, it may be the wing anti-ice system of thevehicle on which the engine 101 is installed.

If it is inappropriate to direct excess capacity anywhere, for exampleif further heating of using anti-ice systems may cause damage given theatmospheric conditions, or the energy storage system is full, then in anembodiment step 1403 solely maximises P_(H) up to P_(D) to assist in thereduction of the high-pressure spool speed.

If the question asked at step 1402 is answered in the negative, to theeffect that the power demand P_(D) is greater than the maximum powergeneration capability of the HP motor-generator 117, P_(Hmax), thencontrol proceeds to step 1404 where first the power generation of the HPmotor-generator 117, P_(H), is maximised, then the power generation ofthe LP motor-generator 119, P_(L), is maximised to supply remainder ofP_(D).

Following optimisation of the power generation strategy to satisfy thepower demand P_(D) in the preceding steps, the fuel flow W_(F) meteredby the fuel metering unit 308 is reduced at step 1206.

In an alternative, embodiment the excess capacity P_(Hmax)-P_(D) may bedirected to the LP motor-generator 119. This may be possible due to thisexcess power representing a small proportion of the power generated bythe low-pressure turbine 108, therefore leading to a very small changein thrust generated by the fan 102. Whilst the change in thrust may besmall, the effect on the high-pressure spool is large in termspreventing an increase in mass flow at the point of reduction of fuelflow, and thereby on the ability to prevent weak extinction.

FIGS. 15A & 15B

When an increase in power demand P_(D) occurs, the power controller 309must respond by in turn demanding an increase in power output by the gasturbine engine.

FIG. 15A illustrates an exemplary increase in power demand P_(D) ofmagnitude dP_(D) within a timeframe dt. FIG. 15B shows a characteristicfor an exemplary axial flow compressor, forming part of a single gasturbine spool coupled to a generator. In order to satisfy the increasein power demand dP_(D) the specific work of the turbine must increase.The steady-state working line is shown at 1501, with the surge lineshown at 1502. To increase the work by the engine, an increase in fuelflow is required.

For the situation in which the generator load follows the step of FIG.15A, the spool may be held at constant corrected speed, or allowed toaccelerate to a higher non-dimensional speed. The movement of theoperating point of the exemplary compressor for each option is shown onthe characteristic of FIG. 15B. Line 1503 shows the movement of theoperating point at constant corrected speed. Line 1504 shows themovement of the operating point to a higher corrected speed. It may beseen that responding in this manner would mean that as the generatorload increases, the compressor non-dimensional speed exhibits a slightinitial reduction as a greater proportion of the turbine work is used todrive the generator rather than the compressor. As fuel flow increases,the compressor operating point moves towards and in both examplesexceeds the surge line 1502.

Thus it may be seen that at low engine throttle settings in particular,such an increase in power demand P_(D) may unchecked cause a compressorto enter surge, requiring additional handling systems prevent this andguarantee adequate surge margin. In practice, this situation may forexample occur in an aircraft engine during descent when anti-ice systemsneed to be enabled but the engines are at an idle setting.

FIGS. 16A & 16B

In the present embodiment, however, the approach is taken to utilise theenergy storage system to mitigate the possibility of surge. Thus, asillustrated in FIG. 16A, the same increase in power demand P_(D) ofmagnitude dP_(D) within a timeframe dt is demanded. Instead of thissolely being met by one or both of the HP motor-generator 117 and the LPmotor-generator 119, it is met during the manoeuvre by the battery 305.Thus, as shown in the Figure, initially the power demand is met by thebattery 305, as shown by the shaded region 1601. As the engine 101accelerates, the proportion provided by the motor-generator(s) increasesgradually until the new power demand is fully met by the engine 101.

FIG. 16B shows the transient working line 1602 on a compressorcharacteristic when this approach is adopted. As the initial increase inpower demand is fulfilled by a different energy source to the core gasturbine engine, there is no attendant drop in compressor non-dimensionalspeed. In addition, the increase in fuel flow may be tempered, so thatthe raise in working line during the transient manoeuvre is not as greatas in the example of FIGS. 15A and 15B. In this way, adequate surgemargin is maintained, potentially allowing a more optimum compressordesign and/or removal of handling systems.

In an alternative embodiment, the battery 305 may provide all of thehigher power demand whilst the engine 101 accelerates to a highercorrected speed, at which point provision of the power demand Po isswitched from the battery to the motor-generator(s) in the engine 101.

FIG. 17

Steps carried out by the optimiser 402 to achieve the functionalitydescribed previously for an increase in power demand P_(D) are set outin FIG. 17.

Following the identification of an increase in power demand within agiven timeframe dt by the differentiator 502 in the classifier module401, the optimiser 402 enters the corresponding optimisation routine atstep 1701. At step 1702, the optimiser 401 evaluates the operatingpoints of the low-pressure compressor 104 and the high-pressurecompressor 105 for the demanded P_(D). In the present embodiment, thismay be achieved using the engine model 404 and knowledge of the currentpower lever angle setting etc. Alternatively, a look-up table or similarmay be used instead.

At step 1702, the current surge margin in the low-pressure compressor104, dR_(L)/R_(L), and the current surge margin in the high-pressurecompressor 105, dR_(H)/R_(H) are evaluated, again using the engine model404 in the present embodiment, or suitable alternatives if required.

At step 1703, the maximum allowable rate of acceleration for each spoolis evaluated given the requirement to maintain adequate surge marginduring the manoeuvre. In the present embodiment, this may be achieved byreferring to the respective acceleration schedules for the spools.

A question is then asked as to whether acceleration of the high-pressurespool and low-pressure spool only will meet the required power demandwithin the demanded timeframe. If not, for example if the new powerdemand is very high or is required in a very short amount of time, thencontrol proceeds to step 1706 where a decision is taken to utilise thebattery 305 (or other energy storage unit such as a capacitor) tosatisfy the demanded P_(D).

Then, or if the question asked at step 1705 was answered in thenegative, the high-pressure and low-pressure spools are accelerated totheir new operating points by increasing the fuel flow metered by thefuel metering unit 308. As described previously, at this point the newpower demand may then be fully met by one or more of the HPmotor-generator 117 and the LP motor-generator 119. The transition maybe gradual, or the battery 305 may solely supply the additional powerdemand dP_(D) until the new operating points are achieved.

FIGS. 18A & 18B

The effect of power transfer from the LP motor-generator 119 to the HPmotor-generator 117 on the operating point of the high-pressurecompressor 105 is shown in FIG. 18A on the compressors' characteristic.The effect on the operating point of the low-pressure compressor 104 isshown on its characteristic in FIG. 18B.

As power is added to the high-pressure spool, the pressure ratio andflow function increase, as shown in FIG. 18A by the transition from aninitial operating point 1801 to a final operating point 1802 at a highernon-dimensional speed on the compressor's working line 1803.

Extraction of power from the low-pressure spool lowers the working lineof the low-pressure compressor 104. Recalling that the low-pressurecompressor rotational speed is fixed relative to the fan 102, atconstant thrust the low-pressure compressor operating point may onlymove on a constant non-dimensional speed line, in this case speed line1804. Due to the increase in flow function in the high-pressurecompressor 105, the low-pressure compressor 104 is unthrottled and soalso sees a raise in flow function. Thus the operating point moves froman initial operating point 1805 to a final operating point 1806 on speedline 1804 away from the surge line 1807.

It will therefore be understood that controlling the degree ofelectrical power generated by one or both of the HP motor-generator 117and the LP motor-generator 119 allows the mass flow rate of the coreflow C to be varied even at fixed thrust settings. Recalling that thebypass ratio of the engine 101 is defined as the ratio of the mass flowrate of the flow B through the bypass duct to the mass flow rate of theflow C through the core gas turbine, this allows the bypass ratio of theengine 101 to be varied. This has particular advantages in terms ofoptimising the jet velocity of the engine 101 for particular airspeeds.

In an embodiment, power transfer may be used to further vary the bypassratio by operating the LP motor-generator 119 as a generator andoperating the HP motor-generator 117 as a motor.

In this way, it will be understood that the engine 101 may operate as avariable-cycle engine.

It will also be seen that transfer of power from the low-pressure spoolto the high-pressure spool is an effective way of increasing surgemargin in both compressors. It should also be noted that thesefundamental effects occur even in the absence of active power transfer:should the power demand P_(D) be greater than or equal to the capabilityof the LP motor-generator 119, then an increase in surge margin is stillachieved by satisfying the power demand P_(D) by maximising low-pressurespool offtake. This is because a greater enthalpy drop is requiredacross the low-pressure turbine 108, which requires a greater mass flow.The greater mass flow through the high-pressure compressor 105, whilstnot as high as with power transfer, still unthrottles the low-pressurecompressor 104 and increases its surge margin. Thus it will beunderstood that this strategy provides a suitable means for increasingsurge margin in the engine 101.

FIG. 19

Steps carried out by the optimiser 402 to increase surge margin aretherefore set out in FIG. 19.

Following the identification of an operating condition in which surgemargin needs to be increased by the classifier module 401, the optimiser402 enters the corresponding optimisation routine at step 1901. Asdescribed previously, operating conditions such as in high cross windsor other unsteady inlet flow phenomena may trigger entry into thisroutine.

At step 1902, a question is asked as to whether the current power demandP_(D) is less than the maximum power rating of the LP motor-generator119, P_(Lmax). If so, then control proceeds to step 1903 where theoptimiser 402 maximises the power generation by the LP motor-generator119, P_(L) to increase surge margin in the low-pressure compressor 104,and transfers any excess electrical power P_(Lmax)-P_(D) to the HPmotor-generator 117 to raise its operating point up its working line,also increasing surge margin.

If the question asked at step 1902 is answered in the negative, to theeffect that the LP motor-generator 119 is not solely capable ofsatisfying the power demand P_(D), then control proceeds to step 1904where the optimiser 402 maximises the power generation by the LPmotor-generator 119, P_(L) to increase surge margin in the low-pressurecompressor 104. Recall that power extraction from the high-pressurespool normally moves the operating point of the high-pressure compressor105 down its normal working line, but that the extraction of power fromthe low-pressure spool normally moves the operating point up its workingline. Thus in step 1904 the optimiser 402 minimises power generation bythe HP motor-generator 117, P_(H) which substantially maintain itsoperating point at around its steady state value, or slightly higher onits working line.

FIGS. 20A & 20B

Whilst surge margin may be increased by the method of FIG. 19, it may insome cases be beneficial to reverse the direction of the power transfersuch that power is transferred from the HP motor-generator 117 to the LPmotor-generator 119. FIG. 20A shows the movement of the operating pointof the high-pressure compressor 105 in this scenario. FIG. 20B shows themovement of the operating point of the low-pressure compressor 104 inthis scenario. The characteristics include lines of constant isentropicefficiency for the compressors. It can be seen that the transfer ofpower from the high-pressure spool to the low-pressure spool may enablean increase in compression efficiency in the engine 101 by moving theoperating points of the compressors into regions of high efficiency.

FIG. 21

Steps carried out by the optimiser 402 to increase compressionefficiency are therefore set out in FIG. 21.

Following the identification of an operating condition in whichcompression efficiency may be increased by the classifier module 401,the optimiser 402 enters the corresponding optimisation routine at step2101. As described previously, operating conditions such as sufficientlysteady inlet flow may permit entry into this routine.

At step 2102, a question is asked as to whether the current power demandP_(D) is less than the maximum power rating of the HP motor-generator117, P_(Hmax). If so, then control proceeds to step 2103 where theoptimiser 402 maximises the power generation by the HP motor-generator117, P_(H) to increase compression efficiency in the high-pressurecompressor 104, and transfers any excess electrical power P_(Hmax)-P_(D)to the LP motor-generator 119 to lower its operating point on itsworking line, also increasing compression efficiency in this example.

If the question asked at step 2102 is answered in the negative, to theeffect that the HP motor-generator 117 is not solely capable ofsatisfying the power demand P_(D), then control proceeds to step 2104where the optimiser 402 maximises the power generation by the HPmotor-generator 117, P_(H) to increase compression efficiency in thehigh-pressure compressor 105. Furthermore, the optimiser 402 minimisespower generation by the LP motor-generator 119, P_(L) to keep thelow-pressure compressor 104 in as high a region of compressionefficiency as possible.

FIG. 22

As described previously, it is also possible to utilise the HPmotor-generator 117 and LP motor-generator 119 to implement speedlimiting. This may provide advantages in terms of safety, by preventingoverspeed conditions or by managing operation around keep-out zones forexample speed ranges where vibration levels are high.

The limiter 503 monitors the shaft speeds N_(H) and N_(L). In anembodiment, the limiter triggers if a mechanical limit is exceed, i.e.purely on the basis of revolutions per minute. Alternatively, thelimiter triggers on the basis of an aerodynamic limit, i.e. a correctedspeed, and so takes temperatures into account. In this way, thebreakdown of flow in the compressors may be prevented.

Thus, following the identification of a limit condition by theclassifier module 401, the optimiser 402 enters the correspondingoptimisation routine at step 2201. At step 2202, a question is asked asto whether the limit is either the low-pressure shaft speed, N_(L)(either mechanical or aerodynamic), or the high-pressure shaft speed,N_(H) (either mechanical or aerodynamic).

If the trigger was low-pressure shaft speed, N_(L), then controlproceeds to step 2203 in which the optimiser 302 maximises the powergeneration by the LP motor-generator 119, P_(L) to decrease thelow-pressure shaft speed. As described previously with respect to otheroptimisation routines, the electrical energy generated may be stored ifcapacity is available in an energy storage system such as battery 305,or alternatively it may be diverted to other systems such as anti-icingsystems or potentially the HP motor-generator 117.

If the trigger was high-pressure shaft speed, N_(H), then controlinstead proceeds to step 2204 in which the optimiser 302 maximises thepower generation by the HP motor-generator 117, P_(H) to decrease thehigh-pressure shaft speed. Again, the power generated thereby may bediverted to storage or to loads.

Various examples have been described, each of which feature variouscombinations of features. It will be appreciated by those skilled in theart that, except where clearly mutually exclusive, any of the featuresmay be employed separately or in combination with any other features andthe invention extends to and includes all combinations andsub-combinations of one or more features described herein.

The invention claimed is:
 1. A gas turbine engine for an aircraft,comprising: a high-pressure (HP) spool comprising an HP compressor and afirst electric machine driven by an HP turbine, the first electricmachine having a first maximum output power; a low-pressure (LP) spoolcomprising an LP compressor and a second electric machine driven by anLP turbine, the second electric machine having a second maximum outputpower; and an engine controller configured to identify a condition tothe effect that the LP turbine is operating in an unchoked regime, and,in response to an electrical power demand being between zero and thefirst maximum output power, only extracting electrical power from thefirst electric machine to meet the electrical power demand.
 2. The gasturbine engine of claim 1, in which the operation of the LP turbine inthe unchoked regime is identified on the basis of an engine inlet flowMach number being below a threshold and a thrust demand being below athreshold.
 3. A method of reducing fuel consumption at low inlet Machnumbers in a gas turbine engine of the type having a high-pressure (HP)spool comprising an HP compressor and a first electric machine driven byan HP turbine, the first electric machine having a first maximum outputpower, and a low-pressure (LP) spool comprising an LP compressor and asecond electric machine driven by an LP turbine, the second electricmachine having a second maximum output power, the method comprising:identifying a condition to the effect that the LP turbine is operatingin an unchoked regime; and in response to an electrical power demandbeing between zero and the first maximum output power, only extractingelectrical power from the first electric machine to meet the electricalpower demand.
 4. The method of claim 3, in which the unchoked regime isidentified on the basis of an engine inlet flow Mach number being belowa threshold and a thrust demand being below a threshold.